'k:/ M--H g0 -i <^ t^ - ^ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED August 19 '+5 as Advance Restricted Beport L5F25 TESTS OF A LINKED DIFFEEIENTIAL FLAP SYSTEM DESIGNED TO MINIMIZE THE REDUCTION IN EFFECTIVE T) IHH!I1M AT. CAUSED BY POWER By Marvin Pitkin and Robert 0. Schade Langley Memorial Aeronautical Latoratcry I-angley Field, Va. vS?:7» - NACA WASHINGTON ^ACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassified. Some of these reports were not tech- ni "ally edited. All have been renroduced without change in order to expedite general distribution. L-^. OOCtJWENTS DePARTMENT Digitized by tine Internet Arcliive in 2011 witli funding from University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation http://www.archive.org/details/testsoflinkeddifOOIang 1 f(p 710H/ NACA ASR No. L5P25 RESTRICTED NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ADVANCE RESTRICTED REPORT TESTS OF A LINKED DIFFERENTIAL FLAP SYSTEM DESIGNED TO MINIMIZE THE REDUCTION IN EpwECTI^'E DIHEDRAL CAUSED 3Y POV/ER By Marvin Pitkin and Robert 0. Schade SUTvlMARY An Investigation has been made in the Langley free- flight turrnel to determine experimentally'- the effects of a linked differential flap system upon the effective- dihedral characteristics of a scale -oowered airplane 10 model. The differential flap system consisted of two individual flaps so linked as to operate differentially from an initial setting when free and designed to create, rolling moments automatically opposing those created by slipstream effects. Tests were made on the Langley free-flight-tunnel balance and on a trim stand that permitted freedom in- roll and yaw. The results of the tests indicate that the negative dihedral changes caused by power may be materially reduced or completely eliminated by use of a differential flap system. Increasing the flap-differential ratio (ratio of upgoing flap deflection to downgoing flap deflection) to a value above unity increased the effectiveness of the differential flap system in opposing the dihedral changes caused by power but diminished the tendency of the flaps to restore themselves to their initial setting of equal deflec- tion. Little effect was observed when the flap- differential ratio was decreased to a value below unity. Differential flap action also Increased the static directional stability. RESTRICTED iJACA ARR Ko. L5?25 INTRODUCTION Tl.e anplicatlon of power in tractor airplanes generally oav.se s a large decrease in the effective dihedral of such airplanes, particularly at low speeds. For airplanes possessing initially small positive values of effective dihedral in gliding flight, power application may lov«rer the effective dihedral to negative values and induce large and lonsatisf actor y degrees of spiral divergence. These adverse effects carjiot he simply elirainatedhy the expediency of increasing the initial amount of geometric dihedral he cause such a change may provide an excessive amount of dihedral with TDower off and lead to poor or unstable oscillatory characteristics in power-off flight or in pov;er-on flight at high speeds. Prom the previous considerations, it is desirable to seek m.eans of avoiding large dihedral changes due to power. One possible solution proposed by Dr. H. S. Ribner of the Langley Memorial Aeronautical Laboratory involves use of a system of lin-lcage whereby the flaps operate differentially from an initial setting so as to create rolling moments automatically opposing those created by slipstream effects. The results of tests of such a system in the Langley free-flight tunnel are reported herein. A powered model, representative of conventional single-radial-engine fighter airplanes, was em.ployed for all tests. Most of the tests were made on a test stand that permitted freedom in roll and yaw. Measx;re- ments of rolling moments were obtained by use of a calibrated-spring system. Necessary force-test data were obtained on the Langley free-flight-tunnel six- component balance. The effects of differential flap action upon the dihedral characteristics of bhe model were studied for various ratios of differential flap movement. In most cases, the tests were made with vertical and horizontal tail surfaces removed, although a brief study was made of the effect of vertical-tail area upon the effective dihedral. NAG A ARR llo . L5P25 5 SYMBOLS The coefficients and symbols are defined as follows; Cl lift coefficient (L/qS) Q-Q drag coefficient (D/qS) Cx longitudinal-force coefficient (X/qS) Cj rolling -moment coefficient (L/qbS) Cjr, rate of change of rolling -moment coefficient with ^ angle of sideslip, per degree {dCi/di^) Gi^ rate of change of rolling-moment coefficient with angle of yaw, per degree (dC|,/d\i/) L force along Z-axis, -nositive when acting upward, pounds; moment about X-axis, positive when it tends to depress right wing, foot-pounds X force along X-axis, positive vifhen acting forward, pounds D force along wind direction, positive when acting rearward, pcands; diameter of propeller, feet Y force along Y-axis, positive when acting to the right, pounds M moment about Y-axis, positive when it tends to raise nose, foot-pounds N moment about Z-axis, positive when it tends to turn nose to right, foot-pounds /I 2 q dynamic pressure, po'onds per square foot (pPV S wing area, square feet c mean aerodynamic chord, feet b , • wing span, feet V airspeed, feet per second i| 1\TACA ARR No. L5P25 p mass density of air, slugs per cubic foot a angle of attack, degrees 'If angle of yaw, degrees p angle of sideslip, degrees 6f flap deflection, degrees 5f„ right-flap deflection, degrees d5-p^,/d5f-j^ flap-differential ratio (ratio of upgoing flap ' deflection to dovmgolng flap deflection) SfTj/Sf-n mean flap-differential ratio over the first 20° of incremental up deflection To thrust disk-loading coefficient (Effective thrust/pV^^a ) V/nD propeller advance ratio n rotational speed, revolutions per second AGj^^ change in lift coefficient caused 137/ flap deflection -■j. S^ vertical-tail area, square feet 5^ rudder deflection, degrees 5-p left-rudder deflection, degrees THEORY OP DIFFERENTIAL FLAP ACTION An important part of the decrease in the effective dihedral parameter Cj^ caused by power is produced by the lateral displacement of the slipstream over the trailing wing as the airr^lane is sideslipped. The lateral center of -oressure of the additional lift induced by the slipstream moves outboard from its original center position and creates a rolling m.oment about the center of gravity of the airplane. The NACA ARR No. L5P25 variation of this rolling moment with sideslip angle is such as to reduce the effective dihedral. On a wing with a flap, the increase in lift of the trailing wing caused by slipstream displacement is accompanied "by an increase in flap hinge moment. The hinge moment of the flap mounted on the leading wing similarly is decreased when the airplane is sideslipped. The flap system tested is shown in figure 1 and is so designed as to utilize the change of flap hinge moments caused by power in order that the dihedral changes caused by power may be reduced. The system consists of two flaps connected by a mechanical linkage. The arrangement of the control rods and singletree is such that upward deflection of one flap pivots the singletree about its fixed center pivot and thus causes downward deflection of the flap on the opposite wing. The central bar, which provides the fulcrum for the differential action, is used to deflect or retract both flaps equally and is extended and locked in flap-down flight. When the apnlied hinge moment on the differential flaps changes because of slipstream displacement, the traillng-wing flap tends to rise and the leading-wing flap tends to fall. As the trailing-wing flap rises, its aerodynamic hinge moments decrease v/hereas those of the leading-wing flap increase. At some differential setting, equilibrium is again obtained. The aileron effect of the differential-flap deflections produces rolling moments that tend to compensate the rolling moments created by slipstream displacement. In addition to direct slipstream-displacement effects, augmentation by the slipstrearfl of the wing-fuselage interference may make an important contribution to the loss in effective dihedral due to power for lov/- wing airplanes. The effect, however, of such a con- tribution upon the flap hinge moments would probably be similar to that due to the direct effects of slip- stream displacement and, consequently, the basic theory would not be greatly altered. The preceding considerations indicate that dif- ferential flaps are fundamentally a linked-aileron system drooped to some Initial downward setting and having an upfloating tendency. Aileron-linkage theory (reference 1) shows that the operating moments of such D NAG A ARR No. L5?25 a system can be red\;ced if the ailerons, starting from equal deflections, are so linked that the upgolng aileron deflects at a progressively .greater rate than the downgcing aileron ^that is, when the values of the flap- differential ratio^ d6fTj/d5f are greater than unity^ . The effectiveness of a differential flap system in producing rolling moments opposing those created by power can therefore be increased by increasing the differential ratio of the system, above unity, because such a change reduces the restoring moments of the system and thus results in greater incremental flap deflections for a gix'^en hinge-moment change induced by power effects. At some differential ratio, the flap system is neutrally balanced and has no tendency to return to the original condition of equllibrl"um. Differential ratios greater than this value create an overbalanced flap system. APPARATUS Wind Tunnel The tests were conducted in the Langley free- f light tunnel; a complete description of the tunnel is given in reference 2. The free-flight-tunnel six- component balance used in the force tests Is described in reference $• Figure 2 shows the test model mounted on the balance strut in a yawed attitude. All force and moment measurements obtained from this balance are with respect to stability axes. The stability axes (see fig. 3) are a system of axes having their origin at the center of gravity of the airplane and in which the Z-axis is in the plane of sjrrnraetry of the airplane and is perpendicular to the relative wind, the X-axis is in the plane of sjrrmnetry and perpendicular to the Z-axis, and the Y-axis Is perpendicular to the plane of SAnnmetry. Trim Stand Most of the tests were m^ade on a trim stand that was so constructed as to allow the model freedom in roll and ja.-w about the stability axes. The construction of the stand is illustrated in the sketch shown as figure ij., MCA ARR No. L5P25 A photograph of the model mounted, on the trim stand is shovvn in figure S . As shown in figure I4., a calibrated spring was attached to the roll-free bearing for the tests to provide for stability in roll and to permit unbalanced rolling moments to be obtained as a function of the angle of bank. The angle of bank was read visually by means of the calibrated indicator card shown in figure 5« Flap deflections were also read directly from an indicator card by means of a pointer rigidly attached to the inboard end of the left-flap segment. (See fig. 5') Mode 1 The model used in the investigation is generally representative of low-wing radial-engine fighter air- planes and corresDonds to a scale model of a kO-foot- 10 span airolane. A three- view drawing of the model is shown as figure 6 and photographs of the model are shown in figure 7- The dimensional characteristics of the full- scale airplane as represented by the -^^^ — scale model tested in the Langley free-flight tunnel are as follows: Propeller: Diameter, feet 11,7 Number of blades 2 Wing: Area, square feet 266.5 Span, feet [j.0 Aspect ratio 5.71 Airfoil section Rhode St. Genese $5 Incidence at root, degrees Dihedral, degrees Sweepback at quarter percent chord line, degrees 3.2 Taper ratio 2:1 Mean aerodynamic chord, inches 83. 90 Root chord, inches IO7.80 Center of gravity: Back of leading edge of root chord, inches . . 32. 9I Below fuselage center line, inches Percent of mean aerodynamic chord 25 8 NAG A ARxR No. L5P25 Flaps : Type , Split, partial span Span, feet 20 Percent wing span 50 Tail: Vertical tail 1 Total area, square feet ,l3.5l^ Percent wing area 5 Vertical tail 2 Total area, square feet 26.68 Percent wing area . 10 Vertical tail 3 Total area, square feet [{.0.0 Percent v>ring area 15 The model was equipped virith a ll|.. 0-inch diameter,, two-blade propeller set at an angle of pitch of 10° at O.T'^ radius and was DOwered by a direct-current controllable-speed electric motor rated 1— horsepower at 12,000 rpm. The propeller was attached to the motor by direct drive, and an electrical tachometer was installed on the motor to permit direct measurements of propeller speed. Right-hand propeller rotation was used for all tests. The layout of the isolated power unit mounted on the roll bracket is shown in figure 8. The model was equipped with partial-span split flaps of 25 percent chord and of total span 50 percent of the wing span. The flaps when locked and not in differential operation were at an initial setting of 1|0 , The right- and left-wing flaps were linked together through a differential linkage located in the fuselage. Details of the flap linkage are shown in the photograph presented in figure 9* This linkage system was designed to permit variation of the flap-differential ratio (ratio of upgoing flap deflection to downgoing flap deflection). This result was accomplished by moving the two end pivots of the singletree rearward with respect to the fixed central pivot. The linkage system was so arranged as to permit maximum differential flap deflections of 57.5° down and 25° up from the Initial flap setting of Ij.0O down. WACA ARR No. L5P25 9 The Rhode St. Genese 35 airfoil section was used on the model wing because of the high maximum lift coef- ficient of this section at the low Reynolds numbers at which the tests were run. The geometric dihedral of the model measured from the lower surface was set at 0° for all tests. No horizontal tail surfaces v/ere used on the model. Three similar vertical tail surfaces of different areas were installed on the model for some tests. Sketches of these tail surfaces are shown in figure 10. TESTS Test Conditions All tests were run at a dynamic pressure of 1.90 pounds per square foot, which corresponds to an airspeed of about 27 miles per hour at standard sea-level conditions and to a test Reynolds number of 172,000 based on the mean aerodynamic chord of O.67 foot. All forces and moments measured in the tests are with respect to the stability axes (fig. 3), which Intersect at a point located at 25 percent mean aerodynamic chord and on the center line (thrust line) of the fuselage. In order to obtain sizeable power effects upon the effective dihedral all power-on tests were made at Tc - 0«96, a value that represented the maximum thrust obtainable from the motor-propeller unit. This value simulated full-scale brake horsepowers ranging from approximately 3000 to 9000 over the high-lift range. Force Tests Force tests were made with power on and with pro- peller off, and with flaps undeflected and deflected J4.O for various angles of attack and yaw. Some tests were made v/ith pov/er on at angles of attack of 1.0° and 13.5° and at yaw angles of 0° and ±10° to determine the effectiveness of the flap system. For these tests the flap on the left wing was locked at 0° and the deflection of the flap on the right wing was varied from 0° to 70° • A complete thrust calibration of the propeller-motor unit was made to determine the model power characteristics 10 MCA ARR No. L5P25 A plot of the results of this calibration is presented as figure 11. Trim -S t and Tests Trim-stand tests were made to determine the effect of freeing the differential flaps on the effective dihedral of the test model. The influence of flap- differential ratio upon the effective-dihedral char- acteristics was also studied. Other tests were nade to determine the effect of vertical-tail area upon the effective dihedral and the influence of differential flap action uDon the directional stability. Test -procedure .- Tiie effective-diliedral charac- teristics of the model with tail surfaces removed were determined as follows: The model was set at various angles of yaw on the test stand and the corresponding trim angles of bank for the propeller-of f and r)Ower-on conditions were noted by visual observation. The values of the angles of bank thus obtained were converted to rolling -moment coef- ficients b;/- means of the roll-spring calibration. The effective-dihedral parameter C^,,-., (oi" -C^^) was then directly determined from a plot of these rolling-moment coefficients against the corresponding angles of yaw. The same procedure for determining the effective-dihedral characteristics of the model with vertical tail surfaces installed was followed except that the model was free in yaw and was trimmed at the different angles of yaw by rudder deflection. Calibration curves were obtained for each flap linkage by measuring the upgoing flap deflection produced by a given dovmgoing deflection on the opposite wing. Representative calibration curves obtained in this manner are shown in figure 12. It should be noted that these curves are nonlinear. This nonlinearity is char- acteristic of the linkage system employed and results in a change of flap-differential ratio dSf-g/dSf-p) with incremental flap deflection. For definlteness , the term "flap-differential ratio" was defined by the mean slope of the differential curves over the first 20° of incre- mental up deflection. This procedure is considered NACA ARR Ko. L5P25 11 sufficient to identify the general effects of altering flap-differential ratio in the teats. Scope of trim-stand tests .- Trim -stand tests were made to determine the dihedral characteristics of the model for the following conditions: (1) Flaps locked at 14.0°; propeller off; a = 1.0°; cl = 0.93 (2) Flaps locked at l+QO; T_ = O.96; a = 1.0° and 13.5°; C^ = 1.4 and^2.7 (3) Flaps free; T„ = O.96; a = 1.0° and 15.5°; Cl = 1.4 and 2.7 Propeller-off tests were not run at 13.5° angle of attack because the model was completely stalled at that angle. The effect of varying the flap-differential ratio between 0.8 and l.Ij. was studied for condition (3). The effect of vertical-tail area upon the power-on (Tc = 0.96) dihedral characteristics was investigated at an angle of attack of 13*5° with flaps locked and with flaps free at a differential ratio of 1.0. RESULTS AND DISCUSSION The results of the tests are given in figures 15 to 25. Lift and drag data obtained from the force tests are given in figures 13 to 15 for various power and flap configurations. These data show that maximum lift coef- ficients comparable with those obtained on full-scale airplanes were obtained for all test conditions. Effective-Dihedral Characteristics Effect of powe r „- The effect of power application upon the effective -dihedral characteristics of the model with flaps locked at 14.0° is shown in figure 16. These data show that application of power increased the negative slope of the curve of rolling-moment against yaw angle from -O.OOO63 to -O.OO205. This change corresponds to reduction of about 7° 1^ effective dihedral and illustrates the usual effect of power upon the dihedral parameter. 12 NAG A ARR No. L5P25 Because of the particialar geometric conflg-uration used In t^.e tests, the test model possessed 2^ negative effective dihedral in the propeller-off condition. The raodel differed therefore from conventional full-scale airplanes, which generally possess moderately large positive effective dihedral in the propeller-off condi- tion. This difference, however, is only of academic interest inasmuch as the reduction in effective dihedral caused by power is an incremental effect that is con- sidered independent of the Initial value of dihedral in the propeller-off condition. The data of figure l6 also indicate little effect of lift coefficient upon the dihedral characteristics of the model under conditions of constant thrust coef- ficient. This phenora.enon is unusual inasmuch as an increase in lift coefficient generally results in an increase in power effects. The increase in effective dihedral generally' associated with increasing angle of attack (lift coefficient) was probably sufficient at high angles of attack to offset the increased power effects caused by the sa:nie increase in angle of attack. The action of the slipstreajn in producing dihedral changes, -nreviously discussed in the section "Theory of Differential Flap Action," appears to be verified by the results of force tests made to determine the flap effectiveness with power on (figs. I7 and l8j. These data show that when the model was yawed to the right ( ^!/ = 10°) the lift increments contributed by the flar) on the right (trailing) wing were considerably increased because of the action of the displaced slipstream. The reverse was true when the model was yawed to the left (^l/ = -10°). Effects of differential flap action .- Trim-stand- test data showing the effect of freeing the differential flaps on the effective dihedral are given in figure 19. Figure 19(a) shows little effect with propeller off; whereas figure 19(b) shows that with power on for a differential ratio of 1.0 (equal up and down flap deflections), the negative dihedral change caused by power at a lift coefficient of 1 .li was reduced by over 80 percent. A similar effect of differential flap action was also attained at a lift coefficient of 2,7- Additional data showing the effect of varying flap-differential ratio are shown in figure 20. The slopes of these curves, which are indicative of the NAG A ARR Ko. L5P25 15 effective-dihedral characteristics, are shown plotted against nean flap-differential ratio in figure 21. The results presented in figure 21 indicate that, although decreasing the differential ratio below unity slightly reduced the efficacy of the differential flaps in opposing dihedral changes due to power, increasing the fiap-differsntial ratio above unity was benificlal. The data show that the adverse effects of power upon the effective dihedral were completely eliminated v/hen a differential ratio of about l.Oij. was lased at a lift coefficient of l.ij. or -when a differential ratio of about 0.96 was used at a lift coefficient of 2.7. Use of ratios greater than these values reversed the effect of power and resulted in positive increases in the effective-dihedral parameter with power application. Slightly larger effects of differential ratio were usually encountered at the high-lift condition (Cl - 2.7). These differential-ratio tests showed that the differential flap system employed In the tests became overbalanced at differential ratios of about l.ij.. ViTLien overbalance occurred, the flaps locked violently against their stops as soon as power was applied, thereby inducing large and abrupt rolling motions. This action occurred for all angles of yaw. The beneficial effect of increasing flap-differential ratio is in agreement with the theory of reference 1. As shown by the data in figure 22, increasing flap- differential ratio generally resulted in greater incre- mental flap deflections at a given angle of yaw as a result of reduced unbalance of the flap system. Effect of vertical-tail area .- Results of tests' made to determine the Influence of vertical-tall area upon the power-on effective-dihedral characteristics of the model are presented in figure 23 and are summarized in ^figure 2I4.. These data show that adding vertical-tail area up to I5 percent of the wing area had little effect. The general tendency of such addi- tions, however, was to Increase the effective dihedral. Directional Stability Characteristics Rudder-deflection data from the yaw-free tests (fig. 25) Indicate that differential flap action con- siderably increased the static directional stability. This increase In stability is attributed to the drag li^ NACA ARR No. L5P25 changes accornpanylng the differential action. As the airplane is yawed, the flap on the trailing wing moves up and reduces the drag of that v;ing, v/hereas the flap on the leading wing moves down and Increases the drag. These drag changes produce stabilizing yawing moments. Although no yaw-free tests v/ere made at differential ratios other than 1.0, it appears reasonable to assume that increasing the differential ratio will increase the stabilizing action of the flaps in yaw because the larger flap deflections encountered will cause greater drag increments . Remarks about Design Dynamic response .- In brief tests made to determine the dynamic response of the flap system to sudden and sharp yawing motions, the results (obtained by visual observation) indicated no appreciable lag of flap deflection with yawing motion. It should be emphasized, however, that the friction in the flap system in the current tests was held to small values, perhaps smaller than those encountered in full-scale designs. Inasmuch as excessive friction in the flap system could cause the controls to "freeze" in a differential attitude (particularly for small degrees of balance) and thus to induce violent rolling m^aneuvers, the designer should attempt to limit the friction in the flap system to as small a value as is practical. Linkage design .- The principle of the differential flap linkage is not necessarily restricted to the mechanical, pin- jointed type of system. Although differential ratios other than -unity are readily obtained with this type of system, differential flaps could also be linked by means of hydraulic, cam, gearing, or electrical systems. Aerodynamic balance .- Because of the severity of the rolling motions caused by an overbalanced flap system, care should be taken in the design of differential flap systemiS to allov/ a safe margin of unbalance.. Further analytical and experimental work with particular refer- ence to the effects of nonlinsarlty of flap load and moment characteristics is required to establish a quantitative design procedure for differential flap systems . NACA ARR No. L5P25 15 CONCLUSIONS The follov;ing conclusions have been drawn from tests of a differential flap system installed in a scale powered airplane model in the Langley free- 10 flight tunnel: 1. A flap system in which the right and left flaps were linked together and were free to deflect differ- entially materially reduced or eliminated negative dihedral changes caused by power application. 2. Increasing the flap-differential ratio of the flap system (ratio of upgoing flap deflection to down- going flap deflection) above unity increased the effec- tiveness of the flap system in opposing dihedral changes caused by power. 5. Increasing the flap-differential ratio of the flap system reduced and eventually reversed the aero- dynamic tendency of the flaps to restore themselves to their initial setting of equal deflection. I|.. Little effect ur)on the effective-dihedral char- acteristics was observed when the flap-differential ratio was decreased to a value below unity. 5. Differential flap action caused an increase in the static directional stability of an airplane. 6. Further analytical and experimental study is required to develop a quantitative design procedure for differential flap systems. Langley Femorial Aeronautical Laboratory National Advisory Comjnittee for Aeronautics Langley Field, Va. 16 MCA ARR No. L5P25 REPSRENCSS 1. Jones, Robert T., and Nerken, Albert I.: The Reduc- tion of Aileron Operating Force by Differential Linkage. NaCA TN No. 586, I956. 2. Shortal, Joseph A., and Osterhout, Clayton J.: Preliminary Stability and Control Tests in the NAGA Free -Flight Wind Tunnel and Correlation with Full-scale Plight Tesos. NACA TN No. 8IO, I9J+I. 5. Shortal, Joseph Ar, and Draper, John W. : Free- Flight-Tunnel Investigation of the Effect of the Fuselage Lp-ngth and the Aspect Ratio and Size of the Vertical Tail on the Li:.teral Stability and Control. NACA ARR No. 3D17, 1945. NACA ARR No. L5F25 Fig. 1 NACA ARR No. L5F25 Fig. 2 o 2 0) I— I (U O E CO bD T3 o 0) o c rH X2 O (D C x: o tj o bo cj O I O .H x; CO I (XI 0) u bD NACA ARR No. L5F25 Fig. 3 d/recHon Wind direc-Z/on NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS F/gure J, - System of s-^Qd///^y axes . Arrows /nc^/ca/e pQSih^e cf/rec-//on~s of mamen-f^ and -forces - NACA ARR No. L5F25 Fig. 4 >5 QJ C (d J C •H ■X3 C 03 -P CO 6 • ■H i-H U (U -.-= c C c :3 o -p T3 -p C C c s ■H -p X) -p (U r^: >J bD o • H rH iH O-tn e 1 (D 0) cu -p S-I ■•-i tM c ;:! >3 S-i M O C ■H ^ Cl Q. cd cd S-i M ho Cm O 1 ■P M o cd j:: •H CL. -p c 1 0) • u CO (U tM 0) (4-1 tj ■H :3 X) NACA ARR No. L5F25 Fig. 9 x; T3 Q) o M CL B 0) OJ to bo +j 03 to ^ (U CD CI- C c^ ^ I ■H bD 1^ •-^ C ^ QJ t^^ t^ I CU (U 't-i 0) <^ tJ O -C Q. (d ki bo O -iJ O x: CL, 1 Oi (U bD NACA ARR No. L5F25 Fig. 10 Fig. 11 NACA ARR No. L5F25 J.^ ID .8 ^^ % I I .2 O -.2 - V \ N V ^- NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ■ 1 1 1 .2 .4 .6 -S /O ^ropaller advance rof/Of l/'/z^O mod&Z propeUer used /n f/iQ po M/ere d - /rtode/ /©s/s . NACA ARR No. L5F25 Fig. 12 § I 4S 4-0 32 Z4 /6 Q O / , o.ae / . -^z ^ , .9S- / /.OO vV V, y /J 3 ' '/ ^ /^ ^ /.20 /a 7a // ■^ A ^/y^ ^ A ^ / . / NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 1 1 1 1 O S /e 24 3Z 40 Upgoing f/ap d(S-f/ec/-/on p d^ j deg F/guns /2.-/^epns^en/a//^? O ^ § \ .3 2 o .6 4 ^ .1 .2 o 4 & /2 /6 20 Angle of affach ^ (Xjdeg f/gure /3-L/ft oncf drag choracfer/sf/c^ ot poi^ered test mocte/ emp/oycd /n test's at c//tTerent/Qj f/ab systems. Prope//er-otf concf/t/on- f-O"- a r^ /. 90 pounds per square /boJ-. ' NACA ARR No. L5F25 Fie-. 14 2i.2 10 /.8 O 1.5 -.1 M o^^ ^N /.2 1 '\ ^ -.a 1 8 Ch r^ -.5 -.6 -.7 -.5 .6 .2 O A^-o y c Y / J I Y / c ^ J / \ ^ < y A / f / \ / ( Y h i r V / / u / ) / S' ( u y / y n ) T NATfOr COMHITTEl ML AD FOR Al VISORY IIONAUTICS 4 6 /a J6 ao /Ing/e of offack ^ oc^ deg F/Qure M.-L/f/ and drag characten5t/C3 of Doyvereo' Yesf /Doae/ empJoved /n :^-sts 'or dz/Terent/af r/op sydrems. 7E=0.96 ; (^^0°; 9^=0 ; q=/.^0 pounds per square Fig. 15 NACA ARR No. L5F-25 Ang/& cf QftacM , cx^ deg /^/gure /5-L//f one/ drag charactensf/cs cf a poM/ered Tesf moofs/ errp/oyeof /n tests Q^ c//^eren/?Q/ f/cp systems ¥br /md ang/es c/= ycuV. Tc^O.SSj dr^^O ; g=/.90 pounds per ^q,uore /do/. ' NACA ARR No. L5F25 Fig. 16 0> 3 \a ^^ <^^US/0/J/900 ^(J^UJOLU-6uf//Oy Fig. 17 NACA ARR No. L5F25 A yO r- NATIONAL ADVISORY COMMITTEE FOft AERONAUTICS ^1 -^ \ n \ 1 1 k \ \ LJ \ 'J V 1 \ n \ \ /^ \ \ U \ \ 1 \ \ n V \ \ <^ ^ \ \ \ n \ V o \ \ \ ^ n^O ^ o LJ \ \ \ 1 \ \ ^ \ X \ \ x\^ ^> [^ \ g 8 U 4 Q o V) ^^ ?o OJ \- / • • ■ NACA ARR No. L5F25 Fig. 18 i^ ^ "^ oi "^ O » * " • u Fig, 19a NACA ARR No. L5F25 o II II "q II II M " onO < > CO ^j) ^ ^U3/D/j^aoD ^uacciou/-&u///c>^ > G^ ^ o 4 I ^^ $ h: I Ms ^ ^ u ^5 5 II 1^4 ^S5 I NACA ARR No. L5F25 Fig. 19b -8 -4 O 4 8 Angle of yay^^ ^ ^ deg (d) 7^=0.86 (unless oiherm^e /ndicaMd). F/gure /S.- Conc/uded. Fig. 20a NACA ARR No. L5F25 \h>Sr .5 ^ PI II r g Q ^O "^o^^/^d^^^oo ^c/t^u/ooc/-^u///o^' NACA ARR No. L5F25 Fig. 20b Q O Fig. 21 NACA ARR No. L5F25 S' I ^ V 8^ ^-:>^y^s>oua>/Cf(y /G^/:>^(///D- s>y^ // o^yy_^ NACA ARR No. L5F25 Fig. 22a 64 36 48 o Vv'^J K) 40 Q <: V, 3? s f e.4 ^ ^ 16 8 -8-4048 Angle of yai^^ ^^ deg (a) c^-I.O''; Cl =/.4. Figure 22. - Effect of f/ap-d/fferenfial rofio u/xdo the trim position of a. fy-ee-f/ooting d/fferifof/cu f/bp system . F/ap5 jom aga/nst stops f.4, /c=0-9^;^-t.90 pounds per square foot Fig. 22b NACA ARR No. L5F25 6^ 56 I ^/6 3 -Q -4 O ^ a Angle of yaiv, (//jdeg (d) a=JJ.S°j Ci=£.7. Figure 21 .- Conc/uded. NACA ARR No. L5F25 Fig. 23 ~/2. -8 -4 Angle cf yoyv , ^ , deg F/gare 23. - Effect cf i^ert/ca/ - 7t2// area on t/ie ro/f/ng- moment choracter/st/cs cf a pcwerecy rest mode/ /n yow. 7c =0.96 ; Q =^.7: q~/.90 pounds per sqaans -fool: Fig. 24 NACA ARR No. L5F25 ^ S) ^ HI o □ \ ^ i h^. < z o z » £^ O I I I I I .^ J I II II ^ 84>| O 1 to ^ r '^ I I I Co I 9- '' js>^&uui>jt)c/ fa/ps>c//p- S)/ii/o3jj-j NACA ARR No. L5F25 Fig. 25 n si Z Ui i| \ , \i \ \ [ ] \ V ^ ^ V 'c^ \ ^ \ ^' 1 \ \, Ill \ \ i\ \ \ \ o • / s \ \ \ K ) \ s \ \ J \^ 1 OJ H H s s^p *■ ^ ? ^ 00 >| X^ 00 I "^ f I 12 ^ II ^ rii Q) <0 III 5lll ''uo//o9/jop jspprj^ UNIVERSITY OF FLOBIDA I ^ • 31262 08104 946 1 UNIVERSE' CF FLORIDA DOCUMENTS DEPARTMENT 1 20 MARSTON SCIE.NCE UBRARY * RO. BOX 117011 GAINESVILLE. FL 32611-7011 USA'